This disclosure relates generally to gas turbine engines and, more particularly, to cooling techniques for airfoil sections of turbine blades of the engine.
In general, gas turbine engines are built around a power core comprising a compressor, a combustor and a turbine, which are arranged in flow series with a forward (upstream) inlet and an aft (downstream) exhaust. The compressor compresses air from the inlet, which is mixed with fuel in the combustor and ignited to produce hot combustion gases. The hot combustion gases drive the turbine section, and are exhausted with the downstream flow.
The turbine drives the compressor via a shaft or a series of coaxially nested shaft spools, each driven at different pressures and speeds. The spools employ a number of stages comprised of alternating rotor blades and stator vanes. The vanes and blades typically have airfoil cross sections, in order to facilitate compression of the incoming air and extraction of rotational energy in the turbine. The blades are secured to the rotor disk through a blade platform.
High combustion temperatures also increase thermal and mechanical loads, particularly on turbine airfoils and associated platforms downstream of the combustor. This reduces service life and reliability, and increases operational costs associated with maintenance and repairs.
The trailing edge of the blade has been cooled by directing air into an internal cavity proximate to the trailing edge. However, the intersection of the trailing edge of the blade to the platform is also an area of highly localized stress due to centrifugal loads.
Accordingly, it is desirable to provide cooling to the blade proximate to the trailing edge of the airfoil without increasing localized stresses beyond their limits.